The present invention relates to spacecraft technology, and more specifically to static buildup of charge on spacecraft. Even more particularly, the present invention relates to methods of apparatuses for safely discharging static charge buildup on sunlit spacecraft, such as earth orbit satellites.
One problem with modern, high power, solar arrays of spacecraft in geosynchronous orbit is that the backside of the solar array is a large dark surface that collects current from a layer of earth atmosphere referred to as the magnetosphere. Front sunlit surfaces of the solar arrays are almost completely covered with a very good insulator, a solar cell cover glass. Currents collected by the large dark area can drive a chassis of the spacecraft thousands of volts negative with respect to the front sunlit surfaces of the solar array, a situation known as inverted gradient charging. This potential distribution, when it occurs to a sufficient degree, leads to arcing and is the presumed cause of several on-orbit string failures on geosynchronous orbit communication satellites.
More specifically, spacecraft charging is the buildup of charge on exposed external surfaces and dielectrics of geosynchronous spacecraft. This surface charging results from the spacecraft encountering a geomagnetic substorm environment, i.e., a plasma with particle energies from 1 to 50 keV. Two types of spacecraft charging occur. The first, called absolute charging, occurs when the ambient environment charges the entire spacecraft uniformly to a potential relative to the surrounding space plasma. The second type, called differential charging, occurs when parts of the spacecraft are charged to different negative potentials relative to each other. Differential charging can cause strong local electric fields, both on spacecraft surfaces and within dielectrics, and can be the cause of potentially damaging arcs on solar arrays.
The buildup of large potentials on the spacecraft relative to the ambient plasma is not, of itself, a serious design concern. It may compromise scientific missions seeking to measure properties of the space environment, but spacecraft systems referenced to structure ground are not affected by a uniformly charged spacecraft. However, differential charging can lead to electrostatic discharges (ESDs) that are known to have caused satellite failures. Two recently launched direct broadcast satellites, Tempo 2 and PanAm Sat 6, have lost over 20% of their solar array power due to charging induced ESD.
Differential charging occurs because spacecraft surfaces are not uniform in their material properties, surfaces are either shaded or sunlit, and/or the ambient fluxes may be anisotropic. These and other effects can produce potential differences between spacecraft surfaces or between spacecraft surfaces and spacecraft ground. When a breakdown threshold is exceeded between surfaces or within dielectrics, an electrostatic discharge can occur.
The transient generated by ESD can couple into the spacecraft electronics and cause upsets ranging from logic switching to complete system failure. The plasma created by the discharge can induce high current arcs on solar arrays and other power system components. Discharges can also cause long-term degradation of exterior surface coatings and enhance contamination of surfaces. Vehicle torquing or wobble can be produced from ESD. The ultimate results are disruptions in spacecraft operation.
Spacecraft charging results when the electron flux to spacecraft surfaces exceeds the sum of the electron flux leaving the surfaces and the ion flux to the surfaces. The incident electrons are from the magnetosphere. The electron flux from the surfaces is due to photoemission in sunlight and is about 1.0-4.0.times.10.sup.-5 A/m.sup.2. This flux is much greater than any electron flux from the magnetosphere in geosynchronous orbit. Therefore, spacecraft charging can only occur in sunlight when there are insulating surfaces on a spacecraft.
The composition and time evolution of the geosynchronous plasma environment are quite complex. It is standard practice to represent the environment in terms of a temperature and density, assuming a Maxwell-Boltzmann distribution. In that characterization, the environment is typically a cold, dense plasma with a "temperature" of about 1 eV and a density of 10.sup.8 particles/m.sup.3. Spacecraft potentials in such an environment are less than a few volts.
During a geomagnetic substorm, changes in the solar wind cause the high-density, low-energy plasma near local midnight to be replaced by a cloud of low-density plasma, (10.sup.6 to 10.sup.7 particles/cm.sup.3) with energies from 1 to 50 keV. It is the electron flux from this environment, less than 10.sup.-5 A/m.sup.2, that can charge spacecraft dielectric surfaces to thousands of volts negative and may result in an electrostatic discharge. The hot plasma cloud diffuses in a few hours but is replaced many times during the life of the storm, which may last a day or longer. For analysis, a "worst case" environment is shown in Table 1.
If the spacecraft is near local noon when the cloud appears, it may never see the hot plasma and will not charge. If the spacecraft is near midnight, it may experience charging and upsets. If the spacecraft is near local evening, as it moves toward midnight it will pass into the diffusing cloud and a more severe charging environment. If the spacecraft is near local dawn, it may be overtaken by the hot plasma. The problem for the spacecraft designer is that each of these environments represents a unique set of plasma conditions as viewed by the spacecraft and results in a markedly different charging history.
Spacecraft charging can be understood in terms of an RC circuit with non-linear current sources. The models of the circuit and current sources are discussed below. Here, we present some general characteristics of how spacecraft surface potentials change during a magnetospheric substorm.
As discussed above, spacecraft potentials change in two ways, absolute charging, where the entire spacecraft and all its surfaces charge with respect to the surrounding magnetosphere, and differential charging, where potential differences develop between spacecraft surfaces. For absolute charging the spacecraft potential changes as a whole, that is, the dielectric surface voltages are "locked" to the ground reference voltage. This type of charging occurs very rapidly (in fractions of a second), typically during eclipse. Differential charging usually occurs slowly (in minutes) and results in one part or surface of the spacecraft being charged to a potential different from those of other parts of the spacecraft. This differential charging can also change the absolute charging level of the spacecraft. This is the usual mechanism for sunlight charging, which consequently occurs slowly. A typical spacecraft charging time history is characterized by rapid absolute charging to .sup..about. 6,000 volts occurring within tens of milliseconds and the differential charging of .sup..about. 2,000 volts between the spacecraft chassis and the solar array cover glass occurring over a period of hours.
The spacecraft configuration is of major importance in determining spacecraft charging behavior. A three-axis-stabilized spacecraft can have a rather large negative structure potential (a few thousand volts) in sunlit charging events. The dominant areas controlling charging in this case are the backs of the solar array wings.
Worst-case environments should be used in predicting spacecraft potentials. The ambient space plasma and the solar extreme ultraviolet (EUV) are the major sources of spacecraft charging currents in the natural environment. The ambient space plasma consists of electrons, protons, and other ions. All of the particles have energies, which are often described by the "temperature" of the plasma. A spacecraft in this environment will accumulate charge until an equilibrium is reached in which the net current is zero. The net current to a surface is the sum of currents due to ambient electrons and ions, secondary electrons, and photoelectrons. The EUV-created photoelectron emissions usually dominate in geosynchronous orbits and prevent the spacecraft potential from being very negative during sunlit portions of the mission.
The density of the plasma also affects spacecraft charging. A "thin," or tenuous, plasma of less than one particle per cubic centimeter will charge the spacecraft and its surfaces more slowly than a "dense" plasma of thousands of particles per cubic centimeter. Additionally, the current due to a thin plasma can be leaked off by partially dielectric surfaces, and steady-state surface and potential differences may not be as great as those in a dense plasma.
Although the photoelectron current due to solar EUV dominates over most of the magnetosphere, in and near geosynchronous orbit during geomagnetic substorms the ambient hot electron current can control and dominate the charging process. Unfortunately the ambient plasma environment at geosynchronous orbit is very difficult to describe. To simplify this description for design purposes, typically only the isotropic currents and Maxwellian temperatures are presented, and these only for electrons and protons. Useful answers can be obtained with this simple representation. For a worst case static-charging analysis the "single Maxwellian" environmental characterization given in Table 1 is recommended.
The values given in table 1 are a 90th percentile single-Maxwellian representation of the environment. If the worst-case analysis shows that spacecraft surface differential potentials are less than 500V there should be no electrostatic discharge problem. If the worst-case analysis shows a possible problem, use of more realistic plasma parameters should be considered.
TABLE 1 Worst-case geosynchronous plasma environment Electron number density, n.sup.c (m.sup.-3) 1.12 .times. 10.sup.6 Electron temperature, T.sub.e (eV) 1.20 .times. 10.sup.4 Ion number density, n.sub.i (m.sup.-3) 2.36 .times. 10.sup.5 Ion temperature, T.sub.i (eV) 2.95 .times. 10.sup.4
The electron current density from a Maxwellian plasma is ##EQU1##
where n.sub.e is the electron density in #/m.sup.3, and T.sub.e, is the electron temperature in electron volts. The ion current, assuming all ions are protons, is given by ##EQU2##
For example, the electron current density from the environment in Table 1 is 3.3.times.10.sup.6 A/m.sup.2.
In the past, spacecraft designers have successfully prevented the chassis of the spacecraft from charging negative relative to the front sunlit surfaces of the solar array by having sunlit conducting surfaces on the satellite body electrically connected to the spacecraft chassis. These sunlit conducting surfaces emit photoelectrons when in the sun, thereby keeping the chassis of the spacecraft closer to plasma ground, i.e., closer in potential to the sunlit front surfaces of the solar array. Unfortunately, a present trend to higher power spacecraft, e.g., earth orbit satellites, means that solar array areas (specifically, the large dark surfaces on the backside of the solar arrays) are larger compared with the front sunlit surfaces of the spacecraft body. Consequently, it has become more difficult for the relatively smaller photo emitting area of the spacecraft body i.e., the sunlit conducting surfaces of the spacecraft body, to emit the excess electrons collected by the large dark surfaces on the backsides of the solar array.
Field effect arrays were invented more than 30 years ago at the Stanford Research Institute. Research is ongoing at SRI, the Naval Research Lab, and other institutions. Currents of over 100 mA are achievable with a 1 mm diameter array.
Previously, an unsuccessful attempt was made to use a field emitter array to control a rocket potential in the ionosphere. Other proposals have been made to use Field emitter array's as electron emitters for tethered satellites. Thus, improvements are needed approaches for discharging static charge collected on the large dark surfaces on the backsides of solar arrays.